The present invention relates to a system for controlling the orbit of a satellite by adequately controlling a plurality of thrusters.
As well known in the art, the orbit of a satellite is affected by various factors and, when effected beyond a certain range which is allowable for a particular mission, a plurality of gas jet thrusters or like suitable propulsion devices mounted on the satellite are activated to correct the orbit. Additionally, the attitude or orientation of a satellite has to be controlled within an allowable range in relation to a particular mission.
Usually, thrusters for orbit control are positioned such that their thrust vectors extend through the center of gravity of a satellite, whereby disturbance torque which acts on the satellite in the event of activation of any of the thrusters is eliminated. In practice, however, it often occurs that the thrust vectors of thrusters do not extend through the center of gravity of a satellite due to errors introduced at the time of production or, even if they do extend through the center of gravity, gas jetted from the thrusters impinges on a part of the satellite such as solar cell panels or a satellite body, producing disturbance torque. Because the disturbance torque produced by thrusters is far greater than solar radiation pressure and other environmental disturbance torque, the accuracy of attitude control during orbit control inevitably deteriorates unless adequate measures are taken.
An implemention for orbit control known in the art consists in firing thrusters continuously all through the orbit control, i.e., jetting gas from thrusters continuously over the whole period between the start and the end of orbit control. Such an implementation, however, has a drawback that at the beginning and end of orbit control there occurs a stepwise fluctuation of disturbance torque due to the jets from thrusters, resulting that the attitude suffers from a substantial transient while an attitude control function absorbs the stepwise disturbance torque.
A satellite orbit control system disclosed in, for example, Japanese Patent Laid-Open Publication (Kokai) No. 58-161699 provides one approach to solve the above-stated problem of the prior art system, i.e., the transient of attitude which occurs at the beginning and end of orbit control. Specifically, the system disclosed sequentially varies the proportion of firing duration of thrusters to the ON-OFF repetition period of thrusters, which is constant. The ratio of firing duration to the period mentioned is varied in a predetermined pattern.
In such an orbit control system, however, the firing duration is set up unchangeably based on presumable conditions of external disturbance and without regard to actual attitudes of a satellite. The system, therefore, cannot cope with unexpected magnitudes of disturbance and achieve the desired object.